This invention generally relates to an airfoil such as is utilized in an axial flow turbine. More particularly, this invention relates to a particular airfoil profile that reduces the stagnation heat transfer coefficient on the airfoil's surface.
Turbine airfoils utilized in axial flow turbines can operate at extreme temperatures. These elevated temperatures can lead to undesired oxidation and degradation of both the airfoil and platforms. For this reason, a cooling system is typically integrated into the airfoil to reduce the transfer of heat to the turbine airfoil. Known cooling systems focus on reducing heat transfer to all surfaces of the turbine airfoil to provide an overall reduction in airfoil metal temperature.
The region of largest heat transfer coefficient is located about the airfoil's stagnation point located on the leading edge of the airfoil. High temperature core gas encountering the leading edge of an airfoil will diverge around a suction and pressure side of the airfoil. Some of the high temperature core gas will impinge on the leading edge. The point on the airfoil where the velocity of the flowing gas approaches zero is the stagnation point. There is a stagnation point at every spanwise position along the leading edge collectively referred to as the stagnation line.
The heat transfer coefficient near the stagnation point of the airfoil is proportional to the local curvature of the airfoil surface. Therefore, the smaller the curvature or larger the radius of the airfoil section's surface, the smaller the heat transfer coefficient, and the lower the temperature along the airfoil. However, increasing the leading edge radius thereby reducing the local curvature about the stagnation point can undesirably affect aerodynamic performance.
Accordingly, it is desirable to develop and design an airfoil that reduces the surface temperatures of the airfoil at the leading edge while minimizing impact to aerodynamic performance.